The present invention relates generally to a system and method of propulsion in a gas turbine engine and, in particular, to a system and method of propulsion in a gas turbine engine which provides continuous pressure rise combustion.
It is well known that typical gas turbine engines are based on the Brayton Cycle, where air is compressed adiabatically, heat is added at constant pressure, the resulting hot gas is expanded in a turbine, and heat is rejected at constant pressure. The energy above that required to drive the compression system is then available for propulsion or other work. Such gas turbine engines generally rely upon deflagrative combustion to burn a fuel/air mixture and produce combustion gas products which travel at relatively slow rates and constant pressure within a combustion chamber. While engines based on the Brayton Cycle have reached a high level of thermodynamic efficiency by steady improvements in component efficiencies and increases in pressure ratio and peak temperature, further improvements are becoming increasingly costly to obtain.
Accordingly, improvements in engine efficiency have been sought by modifying the engine architecture such that the combustion occurs as a detonation in either a continuous or pulsed mode. Most pulse detonation devices employ detonation tubes that are fed with a fuel/air mixture that is subsequently ignited. A combustion pressure wave is then produced, which transitions into a detonation wave (i.e., a fast moving shock wave closely coupled to the reaction zone). The products of combustion follow the detonation wave at the speed of sound relative to the detonation wave and at significantly elevated pressure. Such combustion products then exit through a nozzle to produce thrust. Examples of a pulse detonation engine are disclosed in U.S. Pat. Nos. 5,345,758 to Bussing and 5,901,550 to Bussing et al.
It will be appreciated that the related 12/024,592 application has a similar configuration to the gas generator of the present invention and employs helical channels having a symmetrical configuration. Nevertheless, the 12/024,592 application depends upon suspending a strong detonation wave within the helical channels on board a rotor member or suspending an oblique shock wave followed by a region of supersonic combustion within the helical channels. The design challenges associated with this concept make near-term application unlikely.
Accordingly, it would be desirable for a mechanism to be developed which sustains continuous pressure rise combustion of a fuel-air mixture within a compact device while mitigating the challenges associated with prior approaches. At the same time, a steady surrounding flow field is promoted, gases upstream and downstream of the device are isolated, and a high enthalpy exit flow ready to do work is produced. Further, it would be desirable for such continuous pressure rise combustion system to be adaptable to a gas turbine engine for both aeronautical and industrial applications so as to provide a substitute for a combustor or possibly eliminate the entire core (i.e., a high pressure compressor, combustor, and high pressure turbine).